A gas turbine engine generally includes, in serial flow order, a compressor section, a combustion section, a turbine section and an exhaust section. In operation, air enters an inlet of the compressor section where one or more axial compressors progressively compress the air until it reaches the combustion section. Fuel is mixed with the compressed air and burned within the combustion section to provide combustion gases. The combustion gases are routed from the combustion section to the turbine section. In particular configurations, the turbine section includes, in serial flow order, a high pressure (HP) turbine and a low pressure (LP) turbine. The HP turbine and the LP turbine at least partially define a hot gas path of the gas turbine engine. The combustion gases are then routed out of the hot gas path via the exhaust section.
As the combustion gases flow through the hot gas path, thermal energy is transferred from the combustion gases to various turbine hardware components such as stator vanes, turbine rotor blades, turbine shroud seals and other turbine hardware components. As a result, it is necessary and/or beneficial to cool the various turbine hardware components to meet thermal and/or mechanical performance requirements.
Typically, a cooling medium such as compressed air from the compressor section is routed through various cooling passages or circuits defined within or around the various turbine hardware components. However, undesirably high thermal stresses in the various turbine hardware components may occur due to thermal gradients associated with high combustion gas temperatures and significantly lower cooling medium temperatures. Accordingly, a system for cooling a turbine engine that reduces thermal stresses in the various turbine hardware components, particularly the turbine shroud assemblies, would be welcomed in the technology.